Circular Mass Accelerator

ABSTRACT

A mass accelerator for launching objects, such as a payload, via rotational acceleration is disclosed. The system may comprise a chamber maintained at near vacuum pressure, a motor that rotates a hub attached to a tethered projectile in a circular motion inside the vacuum chamber, accelerating the payload until the payload reaches a desired launch speed. The payload may be released from the tether upon reaching the desired launch speed and may exit the chamber through an exit port that is opened briefly to allow the payload to exit. In various embodiments, the circular mass acceleration system can be used to launch a payload into space orbit. By employing rotational acceleration via a mechanical approach, the acceleration system provides a cost-effective reusable system for launching objects.

CROSS-REFERENCE TO RELATED APPLICATIONS

The present application is a continuation of, and claims the prioritybenefit of, U.S. patent application Ser. No. 15/133,105 filed on Apr.19, 2016. The disclosure of the above-referenced application isincorporated herein in its entirety for all purposes.

FIELD OF THE INVENTION

The present disclosure relates generally to the field of massacceleration systems, and more specifically to systems and methods forlaunching a projectile using centrifugal acceleration.

SUMMARY

This summary is provided to introduce a selection of concepts in asimplified form that are further described in the Detailed Descriptionbelow. This summary is not intended to identify key features oressential features of the claimed subject matter, nor is it intended tobe used as an aid in determining the scope of the claimed subjectmatter.

Various embodiments of the present disclosure may be directed to anacceleration system for launching a projectile. An exemplary system maycomprise a chamber at vacuum pressure, and a tether having a first endportion, a central portion, and a second end portion. The first endportion may be coupled to a projectile. A motor may rotate a hub coupledto the tether in a circular motion inside the chamber, the rotatingcausing the projectile to accelerate until it reaches a desired launchspeed. A release mechanism may release the projectile from the tetheronce it reaches the launch speed. An exit port may allow the projectileto exit from the chamber upon release from the tether.

According to additional exemplary embodiments, the present disclosuremay be directed to an acceleration system for launching a projectile. Anexemplary system may comprise a chamber at vacuum pressure. The chambermay comprise an inverted dome-shaped roof structure comprising aplurality of roof panels under tension. A tether may comprise a firstend coupled to a projectile in proximity to an edge of the chamber and asecond end coupled to a motor at a center of the chamber. The motor maycause the tether and the projectile to rotate in a circular motion untilthe projectile reaches a desired launch speed. A release mechanism mayrelease the projectile from the tether once it reaches the launch speed.An exit port may allow the projectile to exit from the chamber uponrelease from the tether.

According to still further exemplary embodiments, the present disclosuremay be directed to a method for launching a projectile. An exemplarymethod may comprise coupling a projectile to a tether within a vacuumchamber. At least a portion of air within the chamber may be removed tocreate a vacuum condition within the chamber. The tether and the coupledprojectile may be rotated in a circular motion until a speed of theprojectile reaches a desired launch speed. The projectile may bereleased from the tether, and then launched from the chamber out an exitport.

BRIEF DESCRIPTION OF THE DRAWINGS

Embodiments are illustrated by way of example and not by limitation inthe figures of the accompanying drawings, in which like referencesindicate similar elements.

FIG. 1 is a partial cutaway view of a circular mass acceleratorstructure according to various embodiments.

FIG. 2 is a perspective view of a circular mass accelerator structureaccording to various embodiments.

FIG. 3A is a front cross-sectional view of a circular mass acceleratorstructure according to various embodiments.

FIG. 3B is a perspective view of a circular mass accelerator structureat an angle, according to various embodiments.

FIG. 4 is a schematic view of a tether according to various embodiments.

FIG. 5A is a perspective cutaway view of a circular mass acceleratorshowing a launch event according to various embodiments.

FIG. 5B is a perspective cutaway view of a circular mass acceleratorshowing a launch event according to various embodiments.

FIG. 5C is a perspective cutaway view of a circular mass acceleratorshowing a launch event according to various embodiments.

FIG. 6 is a graph showing the relationship between acceleration andchamber diameter as a function of launch velocity according to variousembodiments.

FIG. 7A is a graph of baseline booster phase flight profile showingaltitude as a function of time according to various embodiments.

FIG. 7B is a graph of baseline booster phase flight profile showingvelocity as a function of time according to various embodiments.

FIG. 7C is a graph of baseline booster phase flight profile showing Machnumber as a function of time according to various embodiments.

FIG. 7D is a graph of baseline booster phase flight profile showingflight angle as a function of time according to various embodiments.

FIG. 8 is a schematic diagram of various trajectories for projectilelaunch velocities according to various embodiments.

FIG. 9 is a flow diagram of an exemplary method for launching aprojectile according to various embodiments.

DETAILED DESCRIPTION

The following detailed description includes references to theaccompanying drawings, which form a part of the detailed description.The drawings show illustrations in accordance with example embodiments.These example embodiments, which are also referred to herein as“examples,” are described in enough detail to enable those skilled inthe art to practice the present subject matter. The embodiments can becombined, other embodiments can be utilized, or structural, logical, andother changes can be made without departure from the scope of what isclaimed. The following detailed description is therefore not to be takenin a limiting sense, and the scope is defined by the appended claims andtheir equivalents.

A mass accelerator can be used for many applications, such as launchingprojectiles downrange or for launching a payload into space. Deliveringa projectile downrange requires sufficient energy to provide sufficientvelocity for the projectile under given launch parameters (e.g., mass ofthe projectile, launch angle, etc.) to accurately reach a downrangeposition. In mass acceleration for space applications, one fundamentalproblem of rocket-propelled launch vehicles is to generate sufficientenergy and velocity to place the payload into orbit. Traditional rocketscarry massive quantities of propellant to deliver payloads that areminute fractions of the overall vehicle sizes and weights. All theperformance and risks are built into a precision, often single-usevehicle that must be highly reliable and inherently costly.

While incremental gains have been made in rocket technologies to reducespace launch costs, alternative approaches are necessary to reduce thosecosts and increase launch rates by the orders of magnitude necessary tocreate exponential growth in the space transportation industry. Sincethe beginning of the space program, ground-based non-rocket launchsystems such as rail guns and ram accelerators have been proposed toachieve this; however, all have employed unproven technologies with costprohibitive initial capital investment. In contrast, the presentdisclosure comprises an innovative acceleration system which may achievevery high launch speeds without the need for enormous power generationor massive infrastructure. In various embodiments of the presentdisclosure, an electric or combustion-driven motor or turbine may spin atether and launch vehicle (which may be a projectile or a payload)inside a vacuum chamber until the launch vehicle reaches a desiredlaunch speed. The launch vehicle may then be then released through atangential exit port into its launch trajectory. In various embodimentscomprising a space launch, the acceleration system may launch the launchvehicle such that it ascends above the appreciable atmosphere, afterwhich an onboard rocket may provide the final required velocity fororbital insertion. Because the majority of the energy required to reachorbit may be sourced from ground-based electricity as opposed to complexrocket propulsion, total launch cost may be reduced by at least an orderof magnitude over existing launch systems.

FIG. 1 illustrates a partial cutaway view of the launch system 100according to various embodiments. The launch system 100 may comprise acircular mass accelerator structure 150 comprising a floor 140 and sidewalls 110. In certain embodiments, a diameter of the circular massaccelerator structure 150 may be approximately 100 meters, although oneskilled in the art will readily recognize that the diameter may be moreor less than 100 meters depending on particular applications. The floor140 and side walls 110 may be constructed of any suitable structuralmaterial or combination of structural materials, such as concrete,steel, composite, and the like. As further illustrated in FIG. 2, a roofstructure 155 may span the distance across the side wall 110. In someembodiments, the roof structure 155 may comprise a plurality of thinwall mild steel wedge-shaped panels 145 coupled together to form aconcave dome shape. As illustrated in the cross-sectional view of FIG.3A according to various embodiments, the entirety of the concave domeroof structure 155 may span the circular mass accelerator structure 150without interior structural support. The roof structure 155 may besupported by a circumferential compression ring. In this configuration,the roof structure 155 may be in a tensile state at operating vacuumpressures that avoids the potential of buckling that may be encounteredwith alternative types of freestanding structures.

The circular mass accelerator structure 150 may comprise a sealedchamber that can be placed under at least partial vacuum by evacuatingat least a portion of the air within the circular mass acceleratorstructure 150. In certain embodiments, the pressure within the circularmass accelerator structure 150 may be reduced to about 10⁻² torr,although a pressure of less than about 10⁻¹ torr may be acceptabledepending on particular applications. At pressures less than about 10⁻²torr, outgassing complications may begin to occur from the roof 155 andwalls if constructed from mild steel, such as A-36. If pressuressubstantially less than about 10⁻² are desired, then the roof 155 andwalls may be coated or constructed of different materials suitable tothose pressures.

A concern in vacuum chamber designs is typically the level or quality ofvacuum desired. In the present disclosure, air is evacuated from thecircular mass accelerator structure 150 only for the purpose ofminimizing drag and aerodynamic heating, so a high level vacuum as isordinarily called for in common vacuum chamber designs may not benecessary. The maximum external pressure imposed on the circular massaccelerator structure 150 is about 14.7 psi which is well within thestrength capability of standard mild steel plate.

FIGS. 1 and 2 illustrate a circular mass accelerator structure 150. Acircular shape has been selected for certain embodiments because of theeven distribution of load forces throughout the structure due to themass of the roof structure and atmospheric forces on the roof 155 whenthe circular mass accelerator structure 150 is under vacuum. However,one skilled in the art will recognize that the circular mass acceleratorstructure 150 could instead be any non-circular shape desired, such asoval or rectangular. However, deviations from a circular design mayimpose design constraints that could drive up overall costs.

FIGS. 1 and 3A also illustrate a generally flat floor 140 of thecircular mass accelerator structure 150. In certain embodiments toreduce costs, the floor 140 may be constructed in a similar fashion tothe roof 155 using a dome-shaped tension structure. In those embodimentsin which the domed floor 140 is not inverted (i.e., convex opposite tothe concave roof 155), the two dome shapes of the roof 155 and the floor140 may reduce the interior volume of the circular mass acceleratorstructure 150, thereby reducing the amount of air to be pumped out toachieve the desired level of vacuum.

Although not visible in FIG. 1 or 3A, the circular mass acceleratorstructure 150 may house an electrically-driven or combustion-drivenmotor. A rotating shaft coupled to the motor may extend upwards, and ahub 120 may be coupled to a terminal end of the shaft. A launch vehicletether 125 may be coupled to the hub 120 at one end and to the launchvehicle 105 at an opposite end. A counterweight tether 130 may becoupled at one end to the hub 120 directly opposite the launch vehicletether 125 and to a counterweight 135 at an opposite end. While thelaunch vehicle tether 125 and counterweight tether 130 are described insome embodiments separately, they can actually be portions of a singulartether coupled to the hub 120.

The launch vehicle tether 125 may be subjected to high tensile loadsduring operation. For a baseline launch vehicle 105 of 2000 kgaccelerated to 3,000 m/s, and spinning at 535 RPM, the load on thetether will be 3.14×10⁸ N. Materials advancements during the last twodecades in ultra-high tensile strength materials such as PBO, UHMPE,aramid, Kevlar, carbon fiber, carbon nanotubes, and combinationsthereof, have enabled launch vehicle tether 125 construction atreasonable cost. At 535 RPM, the tip of the launch vehicle tether 125 istravelling at 2,800 m/s and is in tension with 16,000 g of radialacceleration (assuming a launch vehicle tether 125 length of 50 m). Thelaunch vehicle tether 125 may be aerodynamically shaped to minimize dragand aerodynamic heating at supersonic speeds. In addition, asillustrated in FIG. 4, the launch vehicle tether 125 may have a largercross-sectional thickness T₁ at the end coupled to the hub 120 than across-sectional thickness T₂ at the end coupled to the launch vehicle105. The ratio of T₁:T₂ may be derived by a formula to minimize the massof the launch vehicle tether 125 while maximizing tensile strength.

Because of the highly optimized launch vehicle tether 125 design,minimal electrical harnessing is routed through the launch vehicletether 125 to the launch vehicle 105 for power. Just prior to chamberpump down, all hardline connections for power and command and telemetrymay be disconnected, and the launch vehicle 105 and its payload mayswitch to internal power. Henceforth, command and telemetry for thelaunch vehicle 105 and its payload may be transmitted via RF airlink orother suitable communications technology (wired or wireless).

The launch vehicle 105 may have a low-angle conical shape, designed forhypersonic speeds and contain, for example, a payload, fairing/sabot,avionics, thermal protection, and aerodynamic stabilization systems. Thelaunch vehicle 105 may further comprise stabilizing fins at an aft end.The circular mass accelerator structure 150 may be evacuated to a lowpressure environment as described above to reduce drag and aerodynamicheat load on the launch vehicle 105 and tether. Once accelerated to thedesired launch speed, the launch vehicle 105 may be released into theambient environment through a fast-actuating exit port 115, and thelaunch vehicle 105 may ascend on a ballistic trajectory to its desiredaltitude. If so equipped, the launch vehicle's 105 onboard rocketpropulsion system may provide the final energy necessary to achieve adesired orbital altitude and inclination. As opposed to currentmicrosatellite launch vehicles which have 3 or 4 stages totaling 70 to100 ft. in length, various embodiments of the launch vehicle 105 of thepresent disclosure for a similar payload may have only one stage and maybe about 15 ft. in length.

A sequence of events for launching the launch vehicle 105 from thecircular mass accelerator structure 150 according to various embodimentsis illustrated in FIGS. 5A through 5C (the roof 155 is not shown inFIGS. 5A through 5C in order to more clearly illustrate interiorcomponents of the launch system 100). Once a desired velocity isattained (measured, for example, either as a velocity of the launchvehicle 105 or RPM of the launch vehicle tether 125), the launch vehicle105 may be released from the launch vehicle tether 125 at a point wherea tangential vector of the path of the tether 125 aligns with the exitport 115 as illustrated in FIG. 5A. Simultaneously, the counterweight135 may be released from the counterweight tether 130. In FIG. 5B, thelaunch vehicle 105 continues along the tangential vector and enters theexit port 115. The launch vehicle tether 125 and the counterweighttether 130 continue to rotate, and the counterweight 135 moves away fromthe counterweight tether 130. The launch vehicle 105 leaves the exitport 115 in FIG. 5C, completing the launch sequence. The counterweightcontinues to move outward toward the side wall 110.

Though it operates at even higher speeds, the launch vehicle 105 maydraw on the pool of projectile shapes whose aerodynamics and stabilitycharacteristics have been extensively researched in supersonic andhypersonic test facilities. A variety of approaches may be employed toaddress thermal environments, ranging from ceramic heat shields on theSpace Shuttle and ablative carbon phenolics on Galileo to heat sinktungsten nose caps on the X43 Mach 10 Scramjet vehicle and hypersonicprojectile artillery shells to active cooling such German DLR SharpEdged Flight Experiment (SHEFEX) suborbital rocket transpiration coolingexperiments.

The exit port 115 may comprise any system or structure that can provideadequate sealing when the circular mass accelerator structure 150 isunder vacuum and that can be opened at a high rate of speed, oralternatively pierced by the launch vehicle 105. In various embodiments,the exit port 115 may comprise a fast-actuating door or shutter. Inother embodiments, the exit port 115 may comprise one or more sheets ofa polymeric material such as Mylar which may be pierced by the launchvehicle 105 upon launch.

Although not shown in the previous figures, the circular massaccelerator structure 150 may comprise a second exit port directlyopposite the exit port 115 to capture the counterweight 135 that isreleased simultaneously with the launch vehicle 105 to minimize animbalance on the motor at the time of release. The counterweight 135 maybe a solid material, or a liquid such as water.

The launch system 100 described above may be a fixed-base systemdesigned according to various embodiments to launch a payload of about40 kg to 150 km altitude circular orbit, or to deliver the payloaddownrange. In addition to performance targets, the launch system 100 maytypically be designed with a radial acceleration of 5,000-25,000 galthough launch systems may be designed with radial accelerationsgreater than or less than this range in other embodiments. In exemplaryembodiments, the radial acceleration is 15,000-20,000 g. Overall systemoperating parameters according to exemplary embodiments are enumeratedin Table 1.

TABLE 1 Exemplary Launch System Operating Parameters Parameter ValueVehicle Mass 2,000 kg Payload Mass 43 kg Launch Velocity 2,800 m/s M =8.2 (Sea Level) Orbital Altitude 350 km Circular Mass 100 m DiameterAccelerator Size Rotational Speed 535 RPM

The rotational speed, and eventual launch velocity V, of the launchvehicle 105 may be set by the motor. The maximum designed radialacceleration may determine the length of the launch vehicle tether 125,and consequently the interior diameter of the circular mass acceleratorstructure 150. FIG. 6 illustrates the interrelationship between theacceleration and circular mass accelerator structure 150 chamberdiameter as a function of launch velocity. For example, given theexemplary launch system operating parameters in Table 1 and consultingthe graph of FIG. 6, if a maximum radial acceleration of 16,000 g isdesired at a launch velocity of 2,800 m/s then FIG. 6 indicates that afacility radius of approximately 50 m (100 m diameter) would berequired.

A typical sequence of launch events for exemplary embodiments involvingorbital insertion is presented in Table 2.

TABLE 2 Sequence of Events for Exemplary Launch Event Time Pump Down ofCircular L0-60 min Mass Accelerator Structure Spinup of Motor L0-30 minRelease Launch Vehicle L0-3 msec Launch L0 Fairing/Sabot Separation L0 +45 sec First Burn Start L1, L0 + 60 sec First Burn Complete L1 + 128.1sec Second Burn Start L2 Target Orbital Insertion L2 + 11.3 sec

As would be understood by a person of ordinary skill in the art, theflight profile may be divided into two phases: booster phase and upperstage phase. The booster phase shown in FIGS. 7A through 7D may berepresentative of ballistic launch with drag according to variousembodiments. FIG. 7A illustrates that the altitude is an exponentialfunction wherein the rate of increase in altitude decreases with time.The velocity may decrease sharply for the first 5-10 seconds afterlaunch as illustrated in FIG. 7B as aerodynamic drag is imposed on thelaunch vehicle 105 as it leaves the circulator accelerator structure 150and encounters non-vacuum conditions. Once the boosters ignite, the rateof deceleration decreases markedly to a gradual, near linear decline asthe desired orbital insertion velocity is reached. FIG. 7C illustratesthat the launch vehicle starts out well into the hypersonic realm andthe Mach number decreases until an altitude of 40 km is reached (seealso FIG. 7A), at which point the Mach number gradually increases. Theflight angle as illustrates in FIG. 7D steadily decreases as the curvedflight path of the launch vehicle 105 flattens out as the launch vehicle105 approaches orbital insertion. The mission design may be furtherrefined to match the initial conditions for the upper stage first burn,L1, and optimized for upper stage performance to tailor the upper stageflight profile to meet a target orbit.

Mission design for the launch system 100 may also comprise a booster andan upper stage phase, similar to the upper stage of any other launchvehicle. In this case, the booster phase is defined by the launchvehicle 105 flight conditions at the exit of the circular massaccelerator structure 150. The horizontal contribution of launchvelocity V converts directly to orbital velocity at lower launch anglesθ, but at the cost of greater aerodynamic drag. Trajectory optimizationsillustrate an interrelationship between payload mass, launch angle θ(e.g., angle from horizontal), target orbit and launch velocity V. Inaddition to the higher drag, a higher corresponding total heating mayoccur which may require additional cooling or launch vehicle 105 mass toaccommodate thermal management. The upper stage phase may be tailored todesired requirements in the same way as traditional rocket missions. Anexemplary structure at a launch angle θ is depicted in FIG. 3B.

Other contributors to performance are launch altitude and launch sitelatitude/azimuth. Similar to traditional rocket aerodynamics, altitudedictates the prevailing Mach regime of flight which in turn determinesdrag coefficient. Higher altitudes generally reduce drag coefficient anda corresponding reduction in the increase in velocity required from anupper stage. The same factors contribute to lower heating rates. Thelaunch system 100 may also benefit from higher altitudes since thecircular mass accelerator structure 150 may be evacuated to vacuum morereadily. This may reduce power consumption requirements as well as dragand aerodynamic heating during spinup.

FIG. 8 illustrates a schematic diagram of a trajectory for a launchvehicle 105 according to various embodiments. Within the launch system100, the launch vehicle 105 may be rotated at a high speed under vacuumconditions until a desired launch velocity V is obtained. The launchvehicle 105 may then be released and exit the launch system 100 at apre-determined launch angle θ (relative to horizontal). FIG. 8illustrates that the launch vehicle 105 may achieve different downrangedistances D by varying the launch velocity V, all other factors such asthe mass and shape of the launch vehicle 105 and launch angle θremaining constant. For example, at a launch velocity of V₁ a downrangedistance D₁ may be achieved. Increasing the launch velocity to V₂ mayresult in a downrange distance of D₂, and further increasing the launchvelocity to V₃ may result in a downrange distance of D₃. Increasing thelaunch velocity allows for a smaller onboard rocket to be required tolaunch the launch object out of the appreciable atmosphere and achieveorbital insertion. Although not illustrated in FIG. 8, downrangedistances D₁, D₂, and D₃ as well as orbital insertion may be achieved bykeeping the launch velocity V constant and varying the launch angle θ.

FIG. 9 is a flowchart of an exemplary method 900 for launching aprojectile, also sometimes referred to herein as launch vehicle. At step905, a projectile 105 may be coupled to a tether 125 within a vacuumchamber 150. At least a portion of air within the chamber 150 may beremoved at step 910 to create a near vacuum condition within the chamber150. At step 915, the tether 125 and the coupled projectile 105 may berotated in a circular motion until a speed of the projectile 105 reachesa desired launch speed. The projectile 105 may be released from thetether 125 at step 920, and the projectile 105 may be launched from thechamber 150 out an exit port 115.

Various embodiments of the launch system 100 may be used in anyapplication in which a projectile or payload is desired to be launchedto a downrange position or launched into Earth orbit or beyond. Such amass acceleration system can be used in many applications across avariety of industries. In a space launch application, the launch system100 may be used to deliver payloads such as satellites into Earth orbit,or to deliver space vehicles beyond Earth orbit. The launch system 100may also be used in any application where a large amount of force isapplied to a concentrated area. One such example is surface miningoperations where the launch system 100 could essentially be used as acannon to deliver an object into a hillside to break up the rock. Oneskilled in the art will readily recognize that the launch system 100 maybe applied in any situation where a projectile can achieve a desiredfunction.

While the present disclosure has been described in connection with aseries of preferred embodiments, these descriptions are not intended tolimit the scope of the disclosure to the particular forms set forthherein. The above description is illustrative and not restrictive. Manyvariations of the embodiments will become apparent to those of skill inthe art upon review of this disclosure. The scope of this disclosureshould, therefore, be determined not with reference to the abovedescription, but instead should be determined with reference to theappended claims along with their full scope of equivalents. The presentdescriptions are intended to cover such alternatives, modifications, andequivalents as can be included within the spirit and scope of thedisclosure as defined by the appended claims and otherwise appreciatedby one of ordinary skill in the art. In several respects, embodiments ofthe present disclosure can act to close the loopholes in the currentindustry practices in which good business practices and logic arelacking because it is not feasible to implement with current resourcesand tools.

As used herein, the terms “having”, “containing”, “including”,“comprising”, and the like are open ended terms that indicate thepresence of stated elements or features, but do not preclude additionalelements or features. The articles “a”, “an” and “the” are intended toinclude the plural as well as the singular, unless the context clearlyindicates otherwise.

What is claimed is:
 1. A mass accelerator for launching a payload, themass accelerator comprising: an enclosed chamber under at least partialvacuum pressure; a tether having a first end portion, central portion,and second end portion, the first end portion connected to the payload;a motor that rotates a hub attached to the central portion of the tetherin a circular motion inside the chamber, the rotating causing thepayload to accelerate until it reaches a desired launch speed; a releasemechanism that releases the payload from the tether once the payloadreaches the desired launch speed; and an exit port to allow the payloadto exit from the chamber upon release from the tether.
 2. The massaccelerator of claim 1, further comprising: a counterweight mass coupledto the second end portion of the tether.
 3. The mass accelerator ofclaim 2, wherein the counterweight mass is released from the tether atthe same time as the payload.
 4. The mass accelerator of claim 2,further comprising: a second exit port through which the counterweightmass exits from the chamber, the second exit port located directlyopposite to the exit port for the payload.
 5. The system of claim 1,wherein the chamber has a concave roof.
 6. The system of claim 1,wherein the chamber is of a circular shape.
 7. The system of claim 1,wherein the tether is tapered in geometry such that the first endportion and second end portion are of less mass than the centralportion.
 8. The system of claim 1, wherein the payload comprises anonboard rocket to provide additional required velocity for the payloadto enter a space orbit.
 9. An acceleration system for launching apayload, the system comprising: a chamber under at least partial vacuumpressure, the chamber comprising a concave dome-shaped roof structurecomprising a plurality of roof panels under tension; a tether comprisinga first end coupled to a payload in proximity to an edge of the chamberand a second end coupled to a motor at a center of the chamber, themotor causing the tether and the payload to rotate in a circular motionuntil the payload reaches a desired launch speed; a release mechanismthat releases the payload from the tether once the payload reaches thedesired launch speed; and an exit port to allow the payload to exit fromthe chamber upon release from the tether.
 10. The system of claim 9,wherein the vacuum pressure within the chamber is 10⁻¹ torr or less. 11.The system of claim 9, wherein the vacuum pressure is selected to reduceaerodynamic drag and aerodynamic heating of the tether.
 12. The systemof claim 9, wherein the exit port comprises a membrane that is piercedby the payload as the payload exits the chamber.
 13. The system of claim9, wherein the exit port comprises a high-speed shutter that opens toallow the payload to pass through the exit port.
 14. The system of claim9, wherein the payload comprises a vehicle destined for a space orbit.15. The system of claim 9, further comprising: electrical harnessingrouted through the tether to the payload to provide power to thepayload.
 16. A method for launching a payload into space orbit, themethod comprising: coupling a payload to a tether within an enclosedchamber; removing at least a portion of air within the chamber to createat least a partial vacuum condition within the chamber; rotating thetether and the coupled payload in a circular motion until a speed of thepayload reaches a desired launch speed; releasing the payload from thetether; and launching the payload from the chamber via an exit port. 17.The method of claim 16, wherein the removing at least a portion of theair within the chamber comprises removing at least a portion of the airwithin the chamber such that a pressure within the chamber is 10⁻¹ torror less.
 18. The method of claim 16, wherein the rotating the tether andthe coupled payload in a circular motion comprises rotating the tetherand the coupled payload in a circular motion until centrifugal forces onthe payload reach 5,000-20,000 g.
 19. The method of claim 16, furthercomprising simultaneously releasing a counterweight when the payload isreleased from the tether.
 20. The method of claim 16, furthercomprising: igniting a booster onboard the payload, to bring the payloadto a desired orbital insertion velocity.